Peripheral nozzle spinner rocket



July l26, 1960 s. R. cRocKErr PERIPHERAL NOZZLE SPINNER ROCKET I NTOR. SYDNEY R, OCKETT BY ATTORNEYS July 26, 1960 s. R. cRocKETT PERIPHERAL NOZZLE SPINNER ROCKET 2 Sheets-Sheet 2 Filed May 2. 1956 INVENTOR. SYDNEY R. CROCKETT ATTORNEYS United States PERIPHERAL N ZZLE SPINNER ROCKET Sydney R. Crockett, 'China Lake, Calif., assignor to the United States of America as represented by the Secretary of the Navy The invention described herein maybe manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor. l

This invention relates to improvements in jet propelled rockets, and more particularly to improvements involving spin stabilization of such rockets in the launching thereof, and in flight. Y

It has long been recognized that projectiles in ight require stabilization means of some kind.r In general, two types of stabilization have been utilized, namely nn stabilization and spin stabilization. Fin stabilization is generally accomplished by tins attachedY to a projectile in the contemplated direction of ow ofthe airstream (generally parallel to the longitudinal axis of the projectile) and serve to resist movement out of said direction of ow, and such stabilizers are generally not satisfactorily operable in conditions of cross-wind ight or launching. Spin stabilization is effected in'various ways, as by gun-barrel rifling in the case of gun-fired projectiles or by the use of canted ins or vanes or other torque producing means. In the field of jet-propelled rockets, with which the present invention is particularly concerned, spin stabilization is accomplishedby disposing either the jet propulsion nozzles or spin vanes in a direction at an angle to, or canted relative to, the direction of ight of the rocket, and such types of spin stabilization are effective to increase the truen'ess of ight of the rockets. Spin stabilization of rockets is generallymore eifective than n stabilization to decrease the dispersion rates of the rocket designs to which it isy applied.

In the application of spin stabilization torockets a good deal has been learned of the performance requirements of spin-stabilized rockets, and of the effects of different and ideal spin rates at different phases of the iiight of rockets. Thus, it is now known` that one critical phase ofl iiight is the launching phase and that a relatively high rate of spin at this point is effective to prevent undue dispersion of rockets. Secondly, it is knownthat the rate of spin should be proportional to the rate of speed -fr ideal effectiveness in controlling dispersion. Thus, the ideal situation would be to provide a rocket having means effective to develop high rates of torque for providing relaltively high rates of spin at the launching phase of flight and for providing relatively lower torques proportional to the speed of flight of the rocket inthe accelerating stage occurring during burning of the fuel of the rocket. The use of angularly disposed or canted jet nozzles has proven incapable of producing the ideal arrangement discussed hereinabove because any attempt to arrange such nozzles to give high launching spin results in increasingly excessive spin rates as propellant burnining causes continued acceleration, with possible resultant motorv failure caused by breakdown of metallic and other parts.l

atent O Y Y2,946,261 Patentedduly 26, 1960 launching and the iiight phases. VThusrin conventional` rockets with the jet nozzles positioned aft ofthecspin producing iin structure, comparatively large n means are required Yto provide spin inflight because spin is parted by ambient air forces only, andlsuch structures do not lend themselves to high initial spin in launching because of the low initial speed of the rocket and the diculty or impossibility of providing artificial air streams over such outwardly extending surfaces in launch-l mg. Y c

Forwardly disposed nozzles are, per se, not new and are disclosed, for example, in United States Patent No. 2,412,134 issued on December 3, 1946 to C. VL. Eksergian, and in United States Patent No. 2,500,117 issued on March 7, 1950 to E. F. Chandler. However, prior to the present invention no means was known Yto .utilize such forwardly disposed nozzles to effect a benecially variable spin. Thus, the Eksergian patent does not envisage spin stabilization at all and the Chandler patent utilizes a forwardly disposed auxiliary nozzle to impart a progressively increasing spin rate by forcing substantiallyV all' of the gas produced through canted spin vanes, there being no provision for varying the eifectiveness of the gases on the spin vanes. Y

This invention is 'predicated upon the discovery that the Ycombination of forwardly disposednozzles and of rearwardly disposed. canted spin iins of limited extent make it possible to provide airocketmotorwhich has a high torque development characteristic during launching (when used with appropriatelaunching means).Y and which has the inherent capability of reducing the torque producing force upon emergence ofthe rocket motor from the launching means and of then increasing saidr force proportionately as the speed of the rocket increases. Thus, ideally, if propulsion jets are arrangedin the area ofthe center of gravity of an assembled rocket, so as to cause an envelope of highgpressure gases to strike spin vanes fixed to the after end of the rocket body, an advantageous character ofspin variations is developed as an inherent quality of the arrangement. .c Y It is not intended to limit the broad discovery described in the preceding paragraph to the illustrative example depicted in the drawings. However, said example points out auxiliary advantages ofrgreat importanceY which are inherent in the reduction to practice of the basic premises of the presentinvention. Thus, the gas generating section of a rocket according to the principles of this invention may betransported separately fromthe warhead and assembled at a point of use. By virtue of the radial nature of the gas outlets the gas generatoris non-propul siverwhile so transported, which is a safety feature of great importance. Moreover, suchV separate handlingl of components make possible a selective assembly with various types Vof head components, and even a selection of variable jet outlet sizes, since the degree of insertion .of the gas generating portion into the nose portion controls the eiective size of the jet outlet in each case.LY

It is, therefore, an object of thisinvention to provide novel spin-stabilized rocket motor means wherein thev effectiveness of the Vspin producing forces `are`Y variable to provide optimum spin conditions for the var-ious phases of rocket ight.

A further object of this invention is to provid'enovel A still further object of this invention is to provide novel rocket and launcher means wherein the full volume of combustion gas is utilized in the launcher tube to provide high prelaunching, spin-producing torques.

Another object of this inyention is to provide a novel rocket construction wherein spin stabilization is effected entirely by the reaction, of combustionair on spin vanes, ambient air having no.4 direct spin-producing` effect until after burn-out. i

Yet another object of this invention is to provide4 novel rocket motor constructions wherein the gas generating component may be4 s tored and transported separately from the head component and in a relatively sajfe'nonpropulsive condition. A' 'i A still further object: of this invention is to provide ng el and improved rocket motor! constructions where,-` in the gas generating and head c omponfentsl are separable and wherein variation in jet nozzle outlet size may b e eivfzected by the. mode or degree of connection of the said components. A

Still Ianother object of this invention is, to provide novel rocket motor launching and flight methods leading to simplified and improved variable spin-stabilization characteristics in such rocket motors.l i

Yet another object of this invention is to provide novel rocket means wherein forwardly situated. nozzles are utilized for propulsion and wherein the combustion gases released therethrough envelopes the rocket to irnprove its stability, said envelope of gases serving also, varying degree, to impart spin producing torque to saidA rocket. i l 4 Y Other objects and many of the attendant advantages of this invention will be readily appreciated as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings, wherein:

Fig. 1 is a vertical sectional view through a spin-stabilized rocket according to the present invention;

Fig. 2 is a vertical sectionalv View through the propulsion nozzle element of the rocket motor of Fig. 1, it being notable that said element, as illustrated, may be used as a warhead, since space for explosive payload is provided aswell as adapter threads for coupling to an ogival nose piece; i

Fig. 3 is a vertical sectional view through the gas outlet nozzle element of the rocket motor of Fig. A1;

Eig. 4 is a vertical sectional view of rocket launching means for use with the rocket motor of Fig. 1, the rocket being shown schematically and the gas flow in the rocket and in the launcher being diagrammatically indicated by arrows;

Fig. 5 is an elevational view taken from the right side of the launcher shown in Fig. 4, illustrating the open gas-emitting end construction of the launcher tube; and

Figs. 6, 7 and 8 are schematic views of phases of rocket motor flight illustrating the launching phase, the initial slow free iiight phase, and the accelerated flight phase respectively.

Attention is now directed to the drawings, in which like numerals have been appended to like parts throughout. Fig. 1 is illustrative of one embodiment of a rocket motor according to the present invention, and comprises an ogival nose piece 2, having a screw threaded extension for engagement, as at 4, with a jet propulsion nozzle section 6. Nose piece 2 may be arranged with a chamber 8 adapted to be closed by a disc 10 clamped between the nose piece and the propulsion nozzle section, as shown, and this chamber is adapted to hold fuze means, if desired. Propulsion vnozzle section is shown in enlarged detail in Fig. 2 and is of hollow cylindrical conguration, having `an outer wall 11 extending forwardlyinto screw threads 4 for connection to the ogival nase piece 2, as previously described, and a generally rneiially positioned transverse wall 12, dening a chamber 14v which may be loaded with an explosive charge if the performance of a warhead function is desired, or may be used to contain instrumentation for test purposes. The cylindrical wall 11 extends rearwardly of the tr-ansverse wall 12, as shown, and is screw threaded internally for connection, as at 16, with the gas outlet nozzle element 17 of the rocket motor assembly. The wall 11 is externally cylindrical, as shown, but the rearwardly extending portion is formed or cut away to a slightly greater internal diameter to provide an annular passage 18 for propulsion gases, the passage 1,8n flaring at its outlet end into an annular propulsion nozzle 20, as shown, the Passage. being formed between. Said, Gilt-.away mm?. and gas outlet nozzle 17.

Gas outlet nozzle 17, shown in detail in Fig. 3, serves the dual functions of forming one side of passages 18, 20 and of directing gas generated in motor tube 22 to said passages. For these purposes the nozzle 17 is of external dimensions. related to the internal dimensions of the after portion of the propulsion nozzle 6. Thus, the forward portion 24 of nozzle 17 is of screw threaded external diameter and llength to engage the internal screw threads of the wall 11, as, at 1.6, and to abut the transverse wall 12. The forward portion 24. O15 nozzle, 17 is of generally solid metallic construct-ion, having a closed forward portion and is tapped to provide a rearwardly opening chamber 25 and connecting: radial outlet passages 26. Nozzle 1'7 has a Shoulder 29 l formed by the rearward extremity of thickened forward portion 24I and by a rearwardly extending skirt 30, forming. an integral part of said nozzle and servingA as a means fory attachment to motor tube 22. The external diameter ofnozzle 17 is constant from its forward end to a point aft of the outlet passages 26, and the parts arev sopropor-A tioned that said passages 26 open into the annular passage 18 formed in propulsion nozzle section 6, In the area in which the passage 18 is flared toA form jet nozzle 20, the external diameter of the gas outlet nozzle 1] may be increased as at 3 2, if desired, to provide a more ecient restricted ilow path to the jet nozzle 20, and more eflicient flow may beV further eected, itfdesired by rounding the shoulder formed -at the line` of increasing diameter, as at 33.

A motor tube 22 is appropriately secured tQ the skirt 30 as by soldering, and shoulder means 31 may be pro,- vided on gas outlet nozzle 1,7 for positioning and sup? porting the motor tube. Motor tube 22 serves to retain propellant gas producing means and to support spin varies, as shown. Thus, in the illustrative example given herein, a hollow grain 34 of solid propellent material is retained within said tube, with the hollow center thereof in communication with the chamber 245 o f nozzle 17. The after end of motor tube 22 is sealed by a closure cap 36 tted internally Iinto the end of the tube and secured in this position by -any suitable means, asl fpr example by swaging as -at 38. The ylength of grain 34, motor tube 22, skirt 30 of nozzle 17 and the inner skirt of closure cap 36 are such that the grain is clamped against longitudinal movement between said inner skirt and shoulder 29 of the nozzle 17. AIf desired, suitable exible shock absorbing means may be included between the grain and the metallic pieces, .although snel-l details are not illustrated since they are not necessary to the description of the present invention. Igniting means are required for initiating burning of the propellent grain, and one type of such means is illustrated in Pig. l, wherein an igniter 40, comprising a plastic body member containing suitable squib and black powder components is supported within the cap member 36 by means of a post insulatingly passing through and attached te cap 36 (as by ya riveting techniqne). 'I 'he post 42 also. serves as. a pass-through for insulated lead, Wires 14. adapted to be att-ached to a suitable firing circuit- It should be acted that the external eqnslrfatign Qf the rocket thus far described, is auch that an egival nere piece :Quads inta a eYIindrieal parties et einen (the diameter of jet nozzle section '6.) and that aft of the jet nozzle 2t) the diameter is somewhat reduced. It is an important feature of this invention that the spin vane stabilizing means be of noV greater outer diameter than the larger forward diameter of the rocket so that the rocket may be fired from an open-breech cylindrical launching tube of such inner diameter `as to closely accommodate the rocket, thus obviating any substantial loss of gases forwardly of the jet nozzle Z such Aas would occur if an oversize or open arrangement were necessaryto accommodate larger tins. Spin vanes 46, accordingly, are of a radial extent substantially equal to the difference in diameter between jet nozzle section 6 and motor tube 22. A plurality of such vanes 46 are suitable attached, as by welding, to the after end of motor tube 22. Vanes 46 are spaced about the periphery of the motor tube and similarly angularly disposed relative to the longitudinal axis of the rocket whereby gas currents striking the vanes will exert forces tending to impart spin to the rocket. As a cooperating feature of signicant importance, the jet nozzle outlet 20 is ideally situated substantially at the plane of the center of gravity of the rocket motor, as shown best in Fig. 1.

In actual use, the gas generating section of the described rocket motor (comprising gas outlet nozzle 17, motor tube 22` with spin vanes 46, closure cap 36 and igniter 40) will be stored and/or transported separately from the head or heads with which it is intended to be used and that, since passages 26 are radial, the unit will not be propulsive if accidentally ignited. Moreover, this feature of separability between the gas generating section and the rearwardly extending skirt of the head portion makes it possible to vary the effective size of the jet nozzle 20 since variation of the degree of insertion of nozzle 17 into said skirt changes the size of said nozzle and such variation may be readily achieved in various ways, as by shims, or provision of variously designed screw thread lengths, etc.

The rocket described hereinabove, as best shown in Fig. 1 is fired by impressing a voltage across Wires 44 to cause igniter 40 to be operated to produce hot gases and ame to initiate burning of propellant grain 34 with the resultant generation of propulsion gases. Such gases are emitted through chamber 25, radial passages 26, ananular passage 18, and jet nozzle 20 to cause propulsion of the motor. The propulsion gases, emitted from the forward portion of the motor in this manner are effective to form a rearwardly outwardly aring envelope of gases around the after portion of the rocket as shown by G in' Figs. 6 to 8. These gases strike the angular vanes 46 to create spin creating torque, this effect Varying as the gaseous envelope is varied in its concentration in the vane area, as will be more thoroughly explained hereinafter.

Attention is directed to Figs. 4 to 8, which illustrate launching tube means used with the rocket of the present invention and depict the nature of the spin producing gaseous envelope G for different flight phases of the rocket G. Any open-end'ed tube 50 having an internal cylindrical bore slightly larger than the outer diameter of the forward section of rocket R and of the vanes 46 may be used as a launching tube, and, if desired, the after end of tube 50 may be capped by a vented closure 52, with a transverse wall 54 providing gas venting openings as shown in Fig. 5 and a central opening 56 through which lead wires 44 may be threade, if desired. Thearrows G in Fig. 4 show the path of the gases formed in the rocket upon firing in the launcher. The gases formed in the motor tube move forwardly along the only possible egress path and are diverted radially by passages 26 (see Pig. l) and longitudinally rearwardly by passage 18 and jet 20 to cause rocket motion by virtue of reaction to the gas emission. All of the rearwardly directed gases are compelled to pass through the vanes in moving to the vents in the cap 52 for release, because there is no other pos-Y since the speed is relatively slow the spreading is substantial before the vanes are affected thereby, and the spin torque is decreased to an estimated one-sixth of the launching spin-torque. When this occurs the axial thrust is increased and axial velocity increases more rapidly, because less of the forces are being utilized to create spin.

As the rocket thus increases its axial velocity, the gaseous envelope is forced inwardly toward spin vanes 46 by the increasing pressure of the ambient air (see Fig. 8) thus increasing the rate of spin. `This is desirable, since optimum stability requires increased rate of spin with an Yincrease in axial velocity.

From the above it may be seen that the present invention provides new and improved rocket motor construc-A tions wherein ideally variable spin rates are achieved. Moreover, the present invention provides novel rocket launching and flight techniques leading to improved automatically variable spin-stabilization in the launching of rockets and in the free ght thereof.

Only a single structural example illustrative of the Y principles of this invention has been specifically described in this specification. Obviously, many modifications and variations of the present invention are possible in the light of the above teachings. Y

tIt is therefore to be understood that the scope of the invention is to be restricted only by the scope and limitations of the appended claims and not by the details of the said specifically described exemplary modification.

What is claimed is:

1. A spin stabilized rocket adapted to be launched from a launching'tube having a cylindrical bore open at opposite ends thereof, said rocket comprising a nose portion having a maximum diameter slightly less than the diameter of said bore and a portion disposed rearwardly ofthe nose portion of a diameter to provide an annular space between its outer surface and the bore, a plurality of angularly spaced outwardly extending vanes on said second portion disposed angularly to the longitudinal axis of the rocket, the outer peripheries of the vanes being substantially the same diameter as the maximum diameter of said nose portion, whereby the nose portion and vanes prorvide guide means for the rocket to rotate about and translate along. the axis of said bore, and an annular outwardly and rearwardly flaring nozzle disposed forwardly of said vanes adapted to discharge an envelope of gases rearwardly and through said space for impingement against said vanes to impart relatively high initial spin .to the rocket while it is in the tube, said nozzle being constructed and arranged relative to said vanes such that after the rocket leaves the tube the envelope spreads outwardly from the vanes but still partially impinges same, thereby decreasing the spin torque and increasing the elective longitudinal thrust to thereby increase the velocity of the rocket, the envelope being forced inwardly by surrounding ambient air as the rocket further increases in speed to thereby eiect impingement of an increased portion of said envelope on said vanes and produce increased spin torque until the envelope ceases, said vanes serving thereafter to continue spin of the rocket by action of ambient air thereon.

2. A rocket in accordance with claim 1 constructed and arranged to permit adjustment of the size of the throat of said nozzle upon relative longitudinal adjustment of said nose portion and the rearwardly disposed portion, where Btexesces'ited in the f1.1@ of. this patent.

STATES PATENTS 8 Hoagland Dec. 10, 1246 ,.,.,'f-,. -f- Sept' 6: Hickman ----1 Apr. 11, 19.505V Chandler Apr. 18, 1.14950 Chandler Oct. 3,Y 1950 Aflfisano A V Sept. 23,1952

FOREIGN PATENTS 1 Great Britain Jan. 2,l 1,940 

